The present invention relates generally to gas turbine engines, and more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in various turbine stages for powering the compressor and producing working by powering an upstream fan in a typical turbofan aircraft engine application.
The high pressure turbine of the engine includes a turbine nozzle at the outlet end of the combustor which channels the combustion gases into a row of first stage turbine rotor blades mounted to a supporting rotor disk which in turn drives the compressor during operation. The turbine nozzle includes a row of hollow airfoil vanes extending radially between inner and outer bands.
The nozzle is supported in the engine either at its outer band or at its inner band, and is typically formed in circumferential segments for accommodating thermal expansion and contraction as the hot gases are discharged from the combustor and between the nozzle vanes. All of the nozzle components are cooled during operation by using a portion of the pressurized air bled from the compressor in various cooling circuits.
For example, the nozzle vanes are hollow, and compressor discharge air is circulated therein for internal cooling thereof. The sidewalls of the vanes include various rows of film cooling holes through which the spent cooling air is discharged from inside the hollow vanes in thin films of cooling air along the external surfaces of the vanes for providing thermal insulation thereof.
The inner and outer bands of the nozzle are also cooled by the compressor discharge air in additional cooling circuits. However, the bands are relatively thin and are typically cooled by channeling the cooling air over the outboard surfaces thereof, with film cooling holes extending through the bands to the inboard surfaces thereof which confine radially the combustion gas flow therebetween.
The inner and outer bands typically include radially extending flanges which cooperate with adjoining components of the engine for both mounting and sealing the turbine nozzle therewith. Although the flanges are not directly exposed to the hot combustion gases of the turbine flowpath, they provide additional weight and thermal mass which affect performance of the engine.
Weight is the paramount design feature in an aircraft engine and must be minimized for maximizing efficiency of the engine. Thermal mass affects thermal stresses generated during operation, and also affects the durability and life of the turbine nozzle.
A flange configured for sealing a nozzle band to adjacent components may be relatively thin and lightweight, and is relatively easy to cool. However, a flange configured for mounting or supporting the turbine nozzle is relatively thick for carrying the substantial reaction loads through the supporting loadpath, and is therefore more difficult to cool in view of its larger thermal mass and weight.
Accordingly, it is desired to provide a turbine nozzle having an improved supporting flange with selectively reduced thermal mass and weight.